Turbine engine



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May 16, 1967 A. H. BELL. 3,319,931

TURBINE ENGINE NVENTOR. j; HZf/Z /7. ,Belli E A. H. BELL Ill TURBINE ENGINE May 16, 1967v 4 Sheets-Sheet 3 Original Filed July 25 1963 May 15, 1957 A. H. BELL. 3,319,931

TURBINE ENGINE Original Filed July 25 1963 4 Sheets-Sheet 4 United States Patent O 3,319,931 TURBINE ENGINE Albert H. Beli III, Birmingham, Mich., assigner to Chrysler Corporation, Highland Park, Mich., a corporation of Delaware Original application July 25, 1963, Ser. No. 297,538, now Patent No. 3,252,212. Divided and this application Dec. 23, 1965, Ser. No. 516,047

1 Claim. (Cl. 253-78) This invention relates generally to turbine engines and more particularly to the fixed nozzle vanes and wheel blades employed therein.

This application is a division of my copending application Ser. No. 297,538 tiled on July 25, 1963 for Turbine Engine, now Patent No. 3,252,212.

As well known in the art, turbine engines are comprised of a gas generator section consisting primarily of an air compressor, burner portion and a compressor drive turbine wheel which obtains its energy from the gases flowing out of the burner portion and drives the cornpressor. Some turbine engines, depending on their intended use, also include a power turbine whichwhen placed in the path of the flowing gases, downstream of the compressor drive turbine wheel, serves to provide additional work such as by driving a propeller in a turbojet engine or driving some other output shaft for land or water-based vehicles.

The ultimate output power developed by any turbine engine is primarily dependent on the output of the gas generator section which, in turn, is a function of the cornpressor speed of that particular gas generator.

In many instances it has been found that substantial variations in output horsepower exist as between any two turbine engines even though the turbine engines are of the same design and rated output power. Such variations, which in some cases have been in the magnitude of thirty percent (30%) power loss, exhibit themselves to the greatest extent in the range of compressor speeds of seventy-tive percent (75%) to a hundred percent (100%) of the designed or rated maximum compressor speed. Further, these variations occur most frequently in turbine engines wherein the gas ow through the fixed guide vanes or nozzles and compressor drive turbine wheel blades closely approaches sonic velocity as the compressor speed approaches designed maximum speed.

As a consequence of such power variations, it becomes impossible to predict, with any reasonable degree of accuracy, the expected power output of any particular turbine engine. Fulther, since such variations have been found to exist between engines of identical design, it must be concluded that the engine delivering the lesser power is, at least to some degree, of lesser efriciency and therefore undesirable.

g Now it has been discovered that if certain relationships between the elements comprising the turbine engine are considered as being critical and such relationships are established during the process of manufacturing the turbine engine, that the resulting engines will produce a reasonably predictable output horsepower which is consistently at a relatively higher and more eicient value.

Accordingly, a general object of this invention is to provide, in a turbine engine, fixed guide vanes or nozzles and cooperating compressor drive turbine wheel blades having a relationship established therebetween which enables the turbine engine to exhibit an overall higher eiciency.

Another object of this invention is to provide in a turbine engine, fixed guide vanes or nozzles and cooperating compressor drive turbine wheel blades having a relationship therebetween which enables the turbine engine to achieve a predicted output horsepower which is within a reasonable range ofY tolerances.

Still another object of this invention is to provide a method for achieving a particular critical relationship between fixed guide vanes or nozzles and cooperating compressor drive turbine wheel blades of a turbine engine so as to obtain a turbine engine having a predictable horsepower output.

Other objects and advantages of the invention will become apparent when reference is made to the following description and drawings wherein:

FIGURE 1 is a cross-sectional view of a gas turbine engine adapted to propel driving wheels of a land based vehicle;

FIGURE 2 is a graph illustrating typical compressor curves obtained by plotting the ratio of compressor discharge pressure to compressor inlet pressure against the weight-rate lof air flow through the compressor;

FIGURE 3 is a graph illustrating a typical output horsepower curve of a turbine engine;

FIGURE 4 is an enlarged fragmentary portion of the turbine engine of FIGURE l;

FIGURE 5 is a cross-sectional View taken generally on the plane of line 5 5 of FIGURE 4, looking in the direction of the arrows;

FIGURE 6 is a cross-sectional view taken generally on the plane of line 6 6 of FIGURE 5, looking in the direction of the arrows;

FIGURE 7 is a cross-sectional view taken generally on the plane of line 7 7 of FIGURE 5;

FIGURE 8 illustrates a generally tubular conduit of converging cross-sectional area;

FIGURE 9 is a graph illustrating the relationship between the change in velocity of flow through the conduit of FIGURE 8 for a corresponding change in cross-sectional area; and

FIGURES 10 and 11 illustrate, in cross-section, an arrangement for testing nozzle assemblies and turbine wheels, respectively, in accordance with the teachings of this invention.

Referring now in greater detail to the drawings, FIG- URE 1 illustrates a turbine engine 10 adapted for propelling a land-based vehicle having driving wheels 12. The engine 10 is comprised of a housing 14 having an air intake 16 and exhaust orifice 18. A combustion chamber 20, having any suitable fuel distribution means such as fuel distribution ring 22 therein, is located within the housing 14 between the compressor 24 and compressor drive turbine wheel 26. Compressor 24 and compressor drive turbine wheel 26 are connected to each other by means of a shaft 28 which may also be used for driving or in other ways operating selected engine accessories such as the fuel control schematically illustrated at 30.

Compressor 24, as illustrated, is of the radial flow type and is comprised of a generally circular body portion 32 and cylindrical bearing portion 34 with a plurality of generally radially directed compressor vanes 36 therebetween.

Compressor drive turbine wheel 26 is provided with a plurality of circumferentially spaced blades 38 extending radially outwardly from the wheel rim 40. Upstream of turbine wheel 26, the engine 10 is provided with a plurality of radially directed guide vanes or nozzles 42, sometimes referred to as stators, spaced circumferentially about the body portion 44.

As compressor 24 is rotated, air is drawn in through inlet 16, compressed by vanes 36 and directed into the combustion chamber 20 where fuel, supplied in accordance with a predetermined schedule, is -burned so as to heat the gases therein. The gases flow from the burner chamber 20 through nozzles 42 which accelerate and cause the gases to impinge upon the compressor drive turbine wheel blades 38 in -a manner causing rotative motion of turbine wheel 26 which, in turn, drives compressor 24.

The gases, continuing to ow towards exhaust orifice 18, pass between a second set of guide vanes or nozzles 46, similar to nozzles 42, which direct the gases against blades 48 of an output power turbine wheel 50. Turbine wheel 50 may be connected, as by any suitable power transmission means 52, 54 to the ground-engaging driving wheels 12.

FIGURE 2 is a graph illustrating typical compressor curves S6, 58, 60, 62, 64, 66, 68 and 70 obtained by plotting the ratio of compressor discharge pressure, P2, to inlet pressure, P1, against the air flow through the compressor in pounds per second. Each of the curves can be obtained by maintaining the compressor at selected constant speeds while varying the pressure ratio P2/P1. The respective compressor speeds are indicated, for illustrative purposes, near each curve in terms of percentages of designed maximum compressor speed.

The area above and to the left of dash line 72 represents that area in which compressor instability or surge is encountered. Compressor instability or surge refers to a condition sometimes referred to as hunting; that is, those portions of the compressor curve which tend to atten `out and become more horizontal indicate that slight variations in the ratio of P2/P1 result in comparatively large changes in the air iiow which give `rise to compressor instability. Curve 72 is determined generally by connecting together the respective points, on the individual compressor curves, at which compressor surge is first encountered.

In addition to the compressor surge area, as discussed above, a further important consideration remains. That is, the rapid reduction of the numerical value of the ratio of P2/P1 as illustrated by the generally vertically depending portions of each of the compressor curves 56 through 70. Since the compressor inlet pressure, P1, may be assumed to be constant, the conclusion must be that the downstream or compressor discharge pressure, P2, decreases and the rate of P2 decrease, as illustrated by the slope of the depending portions of the compressor curves, is very rapid as compared to an increment of change in the air flow. The reduction in compressor discharge pressure, P2, is caused by the choking effect of the air Within the compressor. Consequently, a line 74 drawn generally through the points of each of the compressor curves wherein the rate of change of the ratio PZ/Pl starts to become rapid defines an area, generally below and to the right of line 74, which might be referred to as the compressor choke area.

Accordingly, with turbine engines, precautions are taken to assure the operation of the compressor within the limits defined by the surge or stall line 72 and choke line 74. In order to achieve this, it has been the practice, generally, to aerodynamically match the compressor drive turbine wheel and its associated nozzle assembly, as a subassembly, to the compressor. This is sometimes referred to as gross matching. In some instances other limits similar to the choke line 74 and the stall line 72 are determined and employed for aerodynamic reasons. (The aircraft industry has also employed additional means for avoiding at least portions of the compressor stall or surge areas by providing compressor bleed valves which at times and in accordance with selected operating parameters, vent some of the compressor air to the atmosphere.) A turbine engine having such a grossmatched compressor, compressor drive turbine wheel and nozzle assembly, so as to operate between the stall line 72 and choke line 74, should exhibit a characteristic horsepower curve 76, obtained generally by plotting output horsepower and the percent of designed maximum compressor speed along a logarithmic Y-axis and an arithmetic X-axis, respectively, of the graph of FIG- URE 3.

However, it has been found that substantial unpredictable variations in output horsepower still exist between gas turbine engines which have been constructed in accordance with the prior art.

For example, with reference to FIGURE 3, it has been found that while one turbine engine might deliver full rated power, another engine of the same design and rating might well develop substantially less horsepower than that predicted. That is, a particular engine might deviate from the predicted mean curve 76, as at some point 80, and after reaching a maximum output as at point 82 (substantially lower than the rated value represented by point 78) steadily decrease to some point 84 which may represent a developed horsepower less than that of point 80.

Further, it has been found that this phenomenon of horsepower variation, or lost horsepower, will at times occur in the same gas turbine engine which had originally developed the full rated horsepower. As was previously stated, the phenomenon of unpredictable horsepower variation occurs most frequently in gas turbine engines wherein the velocity of gas flow through the compressor drive turbine wheel and the compressor turbine nozzle assembly is in the transonic range as the compressor approaches its designed maximum speed.

In the past, various proposals have been tried in unsuccessful attempts to overcome the problem of horsepower variation. For example, the redesigning, modification and/or inspection of such areas as lubrication systems, gear trains, horsepower demands of engine driven accessories, alignment of turbine components and engine air leakages in no -way alleviated the problem of horsepower variation.

Therefore, assuming point 84 to be the lowest horsepower ever achieved, it can be seen that turbine engines, constructed in accordance with the prior art and having the compressor drive turbine wheel and nozzle assembly grossly matched, as a subassembly, to the compressor, can be expected to produce an output horsepower value ranging anywhere between the limits defined by points 78 and 84. In view of this, it becomes evident that the gross aerodynamic matching of the compressor drive turbine wheel and its nozzle assembly, as a subassembly, to the compressor is in itself insufficient for -assuring reasonable attainment of predicted horsepower and that still another inuencing factor, heretofore not apparent to those skilled in the art, remains as an important consideration in the design and construction of the turbine engine.

It has been discovered that such an influencing factor is the relationship between the compressor drive turbine wheel blades 38 and the nozzle or stator vanes 42. More specifically it is the relationship of the total area available for gas flow between the nozzle vanes 42 as compared to the total area available for gas flow rbetween the compressor drive turbine wheel blades 38.

FIGURE 4 illustrates in greater detail the arrangement and construction of the nozzle assembly 43` and compressor drive turbine wheel 26. Preferably, the nozzle or stator vanes 42 are made integrally with a main body portion 86 which is secured to a portion 44 of the general engine housing 14. This may be accomplished as by means of a nut 88 internally threaded so as to cooperate with a threaded portion 90 in axially urging the stator body 86 against the radial surface 92 of a mounting shoulder 94. Radially outwardly, an annular shroud 96 is formed preferably integrally with the nozzle vanes 42. The shroud 96, in cooperation with the outer surface 98 of the stator body 86, defines an annular passage 97 in which the vanes 42 are located for the directional control of the gases flowing therethrough.

The compressor drive turbine wheel 26, secured to shaft 28 for rotation therewith, has its blades 38 generally within the confines of shroud 96. The clearance between the outermost ends 100r of moving compressor drive turbine wheel blades 38 and the inner stationary surface 102 of shroud 96 is kept to a minimum. The outer surface 40 of compressor drive turbine wheel 26 and the inner surface 102 of shroud 96 also dene an annular passage, generally coaxial with passage 97, for the flow of gases therethrough.

The fiow area as between any two adjacent nozzle vanes 42 will =be the smallest cross-sectional area between such adjacent blades. For example, assuming that a plane indicated by line 6 6 of FIGURE 5 were passed through vanes 42 so as to have the cross-sectional area between the vanes a minimum area, the iiow area 103 therebetween could be determined, generally, from the dimensions H1, W1 and B1 as illustrated in FIGURE 6. Similarly, the flow area of the turbine blades 3S could be determined by passing a plane indicated by line 7 7 of FIGURE 5 through points resulting in a minimum area 104. The flow area between such turbine blades would then be determined, generally, yby the dimensions H2, W2, and B2 as illustrated in FIGURE 7. (Because of the relatively small clearances existing between the ends 100 of turbine blades 38 and the inner surface 102 of stationary shroud 96, the height, H2, in FIGURE 7 can be considered as the distance from surface 40 of compressor drive turbine wheel 26 to the inner surface 102 of the shroud '96.)

As previously stated, the gross matching of the nozzle assembly 43 and compressor drive turbine wheel, as a sub-assembly, to the compressor, even though well known in the art, only assures the engine of operation within the limits defined by the stall line 72 and the choke line 74 of FIGURE 2. Such gross matching does not, however, giver any assurance that the engine will attain the maximum horsepower as represented by point 78 of FIGURE 3, but rather that the horsepower attained will be somewhere between the desired point 78 and the low limit of point 84.

It has been discovered that what is still further required, in addition to the gross matching with reference to the compressor, is a critical or selective matching of the nozzle assembly 43 to the compressor drive turbine wheel assembly 26 `based on the respectitve effective ow areas of each.

Referring to FIGURES 4, 5 and 6 it can be seen that if plane 6 6 determines the smallest cross-sectional area between two adjacent nozzle vanes 42 then the crosssectional areas taken at -local points upstream of the plane 6 6 will define progressively large areas generally as the distance between the local area under construction and the minimum area 103 increases. In other words, the progressively larger areas can be considered as defining, generally, the converging conduit 120 of FIGURE 8 wherein the throat area, A1, would be a dimensional equivalent of the minimum area 103- -defined by W1, B1 and H1 of FIGURE 6 and the various local areas, R11, Am, Am, etc., would constitute the increasing cross-sectional areas upstream of the minimum area 103. Further, just as the throat area, AT, will determine the maximum volume rate of air fiow through conduit 120, so will the minimum area 103 between nozzle vanes 42 determine the maximum volume rate of gas ow therethrough for any given condition. Generally, it can be said that the absolute maximum rates in both arrangements will occur when the velocity of air flow through the most restricted areas is sonic.

Assuming that conduit 120 is under a condition of sonic fiow, that is, the velocity of air flow through throat 122 is Mach-one, a relationship can be established between the respective local areas and the velocity of ow through such, areas. For example, an exponential curve, 124 illustrated in FIGURE 9 can be obtained by graphically plotting the ratio of the local cross-sectional areas, A1 to the throat area, AT, against the velocity of fiow at that particular local area.

From an inspection of curve 124, it can be seen that when the ratio of areas, AL/AT, is 1.000 then the velocity of the fiow at the local area is equal to Mach-One. It

should also be observed, however, that the change in local air velocity, A Mach No., is large as compared to a relatively small change in the local cross-sectional area, as reflected by AR, in that range where the ratio of areas, AL/AT, approaches 1.000.

Even though curve 124 of FIGURE 9 has been discussed primarily with reference to the converging conduit of FIGURE 8, the principles involved apply equally well to the nozzle assembly 43. That is, as previously stated, the minimum area 103 determined lby W1, B1 and H1 is the equivalent of the throat area, AT, of throat 122, and the various local cross-sectional areas upstream of the minimum nozzle area 103I are the equivalents of the local cross-sectional areas A11, Am, Am, etc.

Accordingly, it becomes evident that if the size of the minimum nozzle area 103y is calculated to transmit gas therethrough at sonic velocities and at a specific weightrate of flow, that very slight deviations from the calculated size of the minimum nozzle area as illustrated by FIG- URE 9 will result in relatively large reductions in the velocity of air or gas flow through the minimum nozzle area 103. The immediate effect of such velocity losses is a drastic reduction in the energy available for driving the compressor drive turbine wheel 26 and ultimately the power turbine 50.

It has been discovered that in some instances turbine engines, having a nozzle assembly wherein the blade surfaces 108 and 110 (FIGURE 5) were .003 inch further away from each other than the ideal calculated dimension, experienced about a ten percent (10%) power loss. It has also been found that such turbine engines could be made to produce their full rated horsepower if the nozzle assembly and A'compressor drive turbine wheel 26 in the engine were replaced by -a nozzle assembly and compressor drive turbine wheel which were selectively matched to each other.

For example, let it be assumed that it has been empirically determined that a particular engine requires the effective flow area of the nozzle assembly and the effective flow area of the compressor drive turbine wheel to be in an ideal ratio of 1.000 to 1.500, respectively, as determined at a particular pressure differential across each. (The effective or equivalent fiow areas are determined by total flow and velocity experienced through the entire nozzle assembly 43, that is the summation of the effective fiow areas between the individual nozzle vanes, or the entire compressor drive turbine wheel, as the case may be.)

It is well known that dimensional tolerances are required in any manufacturing operation. Therefore, it becomes impossible to randomly choose any one nozzle assembly out of many and similarly choose a compressor drive turbine wheel and have any degree of assurance that the two will have their respective equivalent flow areas in precisely the assumed ideal ratio of 1.000; 1.500. Therefore, the very tolerances necessary for the production of the nozzle assembly and compressor drive turbine wheel cause the respective equivalent flow areas to vary to the degree suflicient to cause ow velocity losses as described with reference to FIGURES 4-9. Such variations resulting from the manufacturing tolerances also exhibit themselves in causing the respective equivalent fiow areas to vary from the assumed ideal ratio of 1.000; 1.500.

However, the discovery of the cause of the problem of variations in and loss of engine power, does not in and of itself suggest a necessarily practical solution of that problem because dimensional manufacturing tolerances cannot be eliminated nor can they be reduced to a degree which results in prohibitive manufacturing costs.

It has been discovered that the solution of the problem resides in the selective matching-of a nozzle assembly to a cooperating compressor drive turbine wheel by means of matching the flow performance of each of them.

For purposes of illustration, let the following be assumed to be the ideal constants:

( 1) ratio of the total nozzle assembly flow area to the total compressor drive turbine wheel iiow area: 1.000/ 1.500;

(2) total nozzle assembly ow area=10.000 sq. inches;

(3) total compressor drive turbine wheel flow area:

15.000 sq. inches;

(4) velocity of gas flow through nozzle assembly at full rated engine power=Mach 0.94;

(5) velocity of gas flow through compressor drive turbine wheel at full rated engine power=Mach one.

Further, let it be assumed that because of the small dimensional differences obtained due to reasonable manufacturing tolerances, that relatively large changes occur (as previously discussed with reference to FIGURES 8 and 9) in the equivalent iiow areas of both the nozzle assembly and the compressor turbine, so as to result in the following:

(6) the effective or equivalent total nozzle assembly ow area varies, as between assemblies, from 9.700 to 10.300 sq. inches and (7) the effective or equivalent total compressor drive turbine flow area varies, as between turbines, from 14.500 to 15.500 sq. inches.

With the above assumptions it can be seen that a nozzle assembly and compressor drive turbine wheel randomly selected from a large group of each could yield the following equivalent flow area combinations having the corresponding indicated equivalent ow area ratios, which would still satisfy the gross matching requirements as between the compressor and the nozzle assembly and compressor drive turbine wheel as a subassembly:

Combinations I and IV result in equivalent flow areas which are very near the assumed ideal area ratios.

It has been discovered that, in gas turbine engines wherein the velocity of motive gas flow through the nozzle assembly or through the compressor drive turbine wheel is in the range of Mach 0.9 to Mach 1.0 as the compressor approaches its designed maximum speed, the numerical equivalent of the actual area ratio may vary in the range of plus or minus 1.0% from the numerical equivalent of the ideal area ratio. Therefore, combinations I and IV would result in the engine developing its full rated horsepower at point 78 of the horsepower curve of FIGURE 3.

However, the same nozzle assemblies combined in reverse order with the compressor drive turbine wheels, as illustrated by combinations II and III, result in equivalent flow area ratios having a substantial deviation from the assumed ideal equivalent iiow area ratio. Combinations II and III, even though satisfying the gross matching requirements O f the compressor, would, nevertheless,

cause the engine to produce an output horsepower much less than the rated output of the engine such as represented by point 84 of the curve of FIGURE 3.

Accordingly, in 4order to insure proper engine performance so as to consistently achieve a reasonably predictable horsepower output it becomes necessary to selectively match the nozzle assembly and compressor drive turbine wheel to each other.

A method of selectively or critically matching the nozzle assembly and compressor drive turbine wheel is disclosed with reference to the apparatus of FIGURE 10 which is comprised of an outer housing 126 provided with a cover 128 at one end thereof and an inlet passage 130 at the other end. Conduit 132, having a valve 134 serially connected therewith, communicates at its one end with a source of pressurized air while its other end serves to retain an orice plate 136 in proper position so as to maintain the orice 138 within the path of the air owing from conduit 132 and into the plenum chamber 140 defined generally by housing 126 and cover 128.

Y Conduits 142 and 144 are provided so as to communicate the pressures on opposite sides of the orifice plate 136 to a differential pressure gage 146.` Conduits 142, 144, orifice 138 and pressure gage 146 provide means for determining the mass rate of flow through the orifice 138 and therefore the entire system. The mass rate of iiow, W, can be determined by the application of the following general compressible mass ow equation:

where:

(1) G=flow per unit of area;

(2) W=mass rate of tiow;

(3) A=crosssectional flow area of orifice 138;

(4) C=the discharge coeiiicient of orifice 138 (a calibrated function of the Reynolds number as well known in the art);

(5) P0=P2=total pressure upstream of orifice 138;

(6) W=molecular weight of the uid owingg,

(7) T 0=total temperature upstream of orifice 138;

(8) g=acceleration due to gravity;

(9) n--the ratio of the specific heat of the fluid at constant pressure, to the specific heat of the liuid at constant volume.

(l0) =universal gas const-ant; and

(l1) r=pressure ratio of downstream static pressure, PS2,

to upstream total pressure, P0.

A testing fixture 152 having an annular passage 154 formed therein is secured in sealing engagement with cover 128. A plurality of guide vanes 156 are provided within the annular passage 154 so as to impart a swirling motion to the air passing therethrough in order to simulate the angle of air impingement experienced by the nozzle vanes in the actual engine. The nozzle assembly 43 t0 be tested is mounted, as illustrated at the exit of the annular passage 154. l

During testing of the nozzle assembly 43 valve 134 is opened to the degree necessary to establish a desired differential, AP, between the plenum chamber pressure, P3, and the barometric pressure, PB. This pressure differential and effective cross-sectional ow area of the nozzle assembly will then determine the velocity of ow through the nozzle assembly.

Since nozzle assembly 43 and orifice 138 are in series with each other, it becomes evident that the mass-rate of ow, W, must be the same for both the nozzle assembly 43 and the orifice 138. Accordingly, once the massrate of ow, W, is determined for orifice 138, the same mass-rate of ow can be employed in the general compressible mass flow equation as applied to the nozzle assembly. The terms of the equation, as applied to the nozzle assembly will have the following meanings:

(1) A=actual cross-sectional ilow area of nozzle assembly 43;

(2) Wzmass-rate of ow through the nozzle assembly (the same as the mass-rate of ow through orifice 138);

(3) P0=P3=pressure upstream of the nozzle assembly;

(4) m=molecular weight of the uid;

(5) T0=total temperature upstream of the nozzle assembly;

(6) C=the discharge coeicient of the nozzle assembly;

(7) g=acceleration due to gravity;

(8) n=the ratio of the specific heat of the gas at constant pressure, Cp, to the specific heat of the gas at constant volume, Cv;

(9) =universal gas constant; and

(10) r=pressure ratio of downstream pressure, PB, to

upstream pressure, P3.

In applying the general equation to the nozzle assembly, the equation may be rearranged into the following form:

It should Ialso be noted that W, 177, To, g, n and E are the same Values as when the general equation was applied to the orifice 138. IFurther, r, is determined from the pressure dierential, Ps-PB. The only values which are not known are C and A. However, the solution of the equation, as rearranged above, yields a value which is the product of C, the discharge coefficient, and A, the actual cross-sectional ow area or, in other words, the effective or equivalent ow area of the nozzle assembly 43 which is precisely the value desired.

Similarly, the compressor drive turbine wheel 26 can be secured at the outlet end of a test xture 158 (FIG- URE 11) which also has an annulus 154 and guide Vanes 156. A shroud 162 located about compressor drive turbine Wheel 26 and secured atop fixture 158 is provided in order to duplicate the coniining effect that the shroud 96 will have within the engine. The remaining portion of the apparatus of FIGURE 11 can be made so as to be identical with that disclosed in FIGURE l0.

The general procedure employed in testing the nozzle assembly is followed in the testing of the compressor drive turbine wheel 26. That is, the valve 134 is opened to the degree necessary to again establish the diierential, AP, between 'P3 and PB.

The mass-rate of flow, W, through orifice 138 is then determined by the application of the general compressible mass flow equation. Once, W, is determined it then becomes possible to solve the same equation, as applied to the compressor drive turbine wheel, for the combined term, CA, (as discussed with reference to the nozzle assembly of FIGURE 10) of the turbine wheel 26.

By employing the above method it then becomes possible to test a plurality of nozzle assemblies, one at a time, and determine and record thereon the equivalent or effective ow area exhibited by that particular nozzle assembly. Likewise, the compressor drive turbine wheels can be tested and the equivalent iow area, so determined, recorded on the individual compressor drive turbine wheels. It then becomes =a matter of simply selectively matching a nozzle assembly and compressor drive turbine wheel, based on their respective equivalent flow areas, which will result in an equivalent ow area ratio Within the critical limits defined on both sides of the ideal equivalent ow area ratio. Such a selectively matched set, when placed within the turbine engine will insure the engine of not only operating within the limits of the choke and stall lines of FIGURE 2 as achieved by gross aerodynamic matching to the compressor, but will further insure the engine of producing an output horsepower at least very closely approaching the designed maximum 10 output horsepower as represented by point 78 of FIG- URE 3.

Further, it has been discovered that as between any two nozzle assemblies or compressor drive turbine wheels, slight dilerences may exist between the respective curves defining the mass-rate of gas ow therethrough as a function of the pressure differential across such nozzle assemblies or compressor drive turbine wheels. This is believed due to the slightly varying aerodynamic characteristics of the particular nozzle assembly or compressor drive turbine wheel under consideration. Accordingly, the method for determining and critically selecting a particular nozzle assembly to a cooperating compressor drive turbine Wheel can be modified so as to even avoid such slight discrepancies arising out of such varying aerodynamic characteristics.

That is, preferably, when the ideal equivalent ow areas are empirically determined, they should be determined by employing a pressure differential substantially equivalent to that pressure diierential which the nozzle assembly or the compressor drive turbine Wheel, as the case may be, will experience within the engine as during full rated power oper-ation or maximum designed compressor speed.

Subsequently, the pressure diierential employed during calibration of nozzle assemblies would vary from that pressure differential employed for Calibrating compressor drive turbine wheels. For example, with nozzle assemblies, a sucient pressure differential may be indicated when pressure gage indicates a value which is 0.75 times the then existing barometric pressure, whereas, in the case of the compressor drive turbine wheel a reading of 1.17 times the then existing barometric pressure may indicate a proper pressure differential.

In each event, however, the calibration is conducted employing a pressure dilerential substantially equivalent to the pressure differential actually experienced by the test piece under conditions of actual engine operation thereby avoiding any influencing factors which might otherwise exhibit themselves due to the aerodynamic characteristics of the compressor drive turbine wheel or nozzle assembly.

The drawings rand the foregoing specification constitute a description of the invention in such terms as to enable any person skilled in the art to practice the invention, the scope of which is indicated by the appended claims.

I claim:

In a gas turbine engine having a compressor; a compressor drive turbine wheel; and a turbine wheel inlet nozzle assembly; wherein said turbine wheel and nozzle assembly define, each by means of a plurality of radially directed circumferentially spaced blades, turbine wheel and nozzle motive gas flow areas; wherein said nozzle motive gas ow area is less than said turbine wheel motive gas ow area; and wherein the velocity of motive gas flow through at least one of said ow areas is in the range of Mach 0.9 to 1.0 during periods of maximum engine speed operation; the improvement of having said turbine wheel and said nozzle assembly critically matched to each other so that the ratio of said turbine wheel motive gas ow area to said nozzle motive gas ilow area is within one percent of that ratio between said turbine wheel motive gas flow area and said nozzle motive gas ow area which enables said turbine engine to produce its maximum power.

References Cited by the Examiner UNITED STATES PATENTS 2,796,658 6/1957 Aller 29-407 X 3,077,074 2/1963 Collman et al. 60-39.16 X 3,142,475 7/ 1964 Bobo et al. 253-78 FOREIGN PATENTS 785,419 10/ 1957 Great Britain.

JULIUS E. WEST, Primary Examiner.

UNITED STATES PATENT OFFICE CERTIFICATE 0E CORRECTION Patent No. 3,319,931 May 16, 1967 Albert H. Bell III rtified that error appears in the above numbered patlt is hereby ce at the said Letters Patent should read as ent requiring correction and th corrected below.

Column 5, line 47, for "construction" read Consideration line 53, for "RLl" read ALl Signed and sealed this 3rd day of December 1968.

(SEAL) Attest:

EDWARD J. BRENNER Edward M. Fletcher, Jr.

Commissioner of Patents Attesting Officer 

